The present invention relates to a tip turbine engine, and more particularly to an assembly for structurally supporting the compressor rotor and compressor case.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan and a low pressure compressor, a middle core engine, and an aft low pressure turbine all located along a common central axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a central high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor and ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine which rotatably drives the high pressure compressor through the central high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine which rotatably drives the bypass fan and low pressure compressor through a central low spool shaft.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship and rotating central shafts require that several engine cases on the outer portion of the engine directly bear the loads of engine components such as the compressor case.
A recent development in gas turbine engines is the more longitudinally compact tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan. The axial compressor and bypass fan share a common rotor for co-rotation. The common rotor is supported on a front end by a front support that is fixed to a housing via a first set of radially extending struts. The common rotor is supported on a rear end by a rear support that is fixed to the housing via a second set of radially extending struts.
The bypass fan of the tip turbine engine includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor located radially outward from the fan. The combustor ignites the fuel mixture to form a high energy gas stream which drives turbine blades that are integrated onto the tips of the hollow bypass fan blades for rotation therewith as disclosed in U.S. Patent Application Publication Nos.: 2003192303; 20030192304; and 20040025490. The integrated bypass fan-turbine drives the axial compressor through the common rotor. Such an architecture, however, depends on two sets of engine support planes, the first and second radial struts, to support the common rotor. Utilizing two engine support planes may complicate the assembly and may be unnecessary to support the length of the longitudinally compact engine.
Accordingly, it is desirable to provide a load bearing support structure from a single support plane for the compressor case and compressor rotor.